The present invention relates generally to gas turbine engines, and, more specifically, to turbine airfoil cooling.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The hot gases are channeled through various stages of a turbine which extract energy therefrom for powering the compressor and producing work, such as powering an upstream fan in a typical aircraft turbofan engine application.
The turbine stages include stationary turbine nozzles having a row of hollow vanes which channel the combustion gases into a corresponding row of rotor blades extending radially outwardly from a supporting rotor disk. The vanes and blades have corresponding airfoil configurations, and are hollow with corresponding cooling circuits therein. Since the heat loads from the combustion gases vary over the pressure and suction sides of the vanes and blades, the cooling circuits therein are correspondingly different for best using the limited cooling air available.
The cooling air is typically compressor discharge air which is diverted from the combustion process and therefore decreases overall efficiency of the engine. The amount of cooling air must be minimized for maximizing the efficiency of the engine, but sufficient cooling air must nevertheless be used for adequately cooling the turbine airfoils for maximizing their useful life during operation.
Each airfoil includes a generally concave pressure sidewall, and an opposite, generally convex suction sidewall extending longitudinally or radially in span and axially in chord between leading and trailing edges. For a turbine blade, the airfoil span extends from a root at the radially inner platform to a radially outer tip spaced from a surrounding turbine shroud. For a turbine vane, the airfoil extends from a root integral with a radially inner band to a radially outer tip integral with an outer band.
Each turbine airfoil also initially increases in thickness aft of the leading edge and then decreases in thickness to a relatively thin or sharp trailing edge where the pressure and suction sidewalls join together. The wider portion of the airfoil has sufficient internal space for accommodating various forms of internal cooling circuits and turbulators for enhancing heat transfer cooling inside the airfoil, whereas the relatively thin trailing edge has correspondingly limited internal cooling space.
Each airfoil typically includes various rows of film cooling holes extending through the sidewalls thereof which discharge the spent cooling air from the internal circuits. The film cooling holes are typically inclined in the aft direction toward the trailing edge and create a thin film of cooling air over the external surface of the airfoil that provides a thermally insulating air blanket for additional protection against the hot combustion gases which flow over the airfoil surfaces during operation.
The thin trailing edge is typically protected by a row of trailing edge cooling slots which breach the pressure sidewall immediately upstream of the trailing edge for discharging film cooling air thereover. Each trailing edge outlet slot has an exposed outlet aperture in the pressure side which begins at a breakout lip and is bounded in the radial direction by exposed lands at the aft ends of axial partitions which define the outlet slots.
The axial partitions are integrally formed with the pressure and suction sides of the airfoil and themselves must be cooled by the air discharged through the outlet slots defined thereby. The partitions typically converge in the aft direction toward the trailing edge so that the outlet slots diverge toward the trailing edge with a shallow divergence angle that promotes diffusion of the discharged cooling air with little if any flow separation from the sides of the partitions.
Aerodynamic and cooling performance of the trailing edge outlet slots is directly related to the specific configuration of the outlet slots and the intervening partitions. The flow area of the outlet slots regulates the flow of cooling air discharged through the slots, and the geometry of the slots affects cooling performance thereof.
The row of trailing edge outlet slots typically extends for the full radial span of the trailing edge for providing cooling completely therealong. The divergence or diffusion angle of the outlet slots is typically limited to about 7.5 degrees for maximizing diffusion efficiency without effecting undesirable flow separation of the discharged cooling air which would degrade performance and cooling effectiveness of the discharged air.
Accordingly, these geometrical constraints on discharge flow area and divergence angle determine the number and pitch spacing of the outlet slots along the trailing edge.
The portions of the thin trailing edge directly under the individual outlet slots are effectively cooled by the discharged cooling air, with the discharged air also being distributed over the intervening exposed lands at the aft end of the partitions. However, those lands are nevertheless solid portions of the pressure sidewall integrally formed with the suction sidewall and must rely for cooling on the air discharged from the adjacent trailing edge outlets.
Notwithstanding the small size of the these outlet lands and the substantial cooling performance of the trailing edge outlet slots, the thin trailing edges of turbine airfoils nevertheless typically limit the life of those airfoils due to the high operating temperature thereof in the hostile environment of a gas turbine engine.
Accordingly, it is desired to provide a turbine airfoil having improved trailing edge cooling for improving airfoil durability and engine performance.